A component (1, 2) for supporting at least one bearing (3) for a turbine engine (10) comprising: two coaxial walls, internal (4) and external (5) walls respectively, defining a gas flow vein (6) between them and interconnected by a row of arms (7); an external ferrule (50) comprising an internal peripheral edge (51) connected to the external wall (5) and an external peripheral edge (52) connected to an external mounting flange (53); an internal ferrule (40) comprising an external peripheral edge (41) connected to the internal wall (4) and an internal peripheral edge (42) comprising an internal mounting flange (43); at least one of the ferrules (4, 5), which at the peripheral edge (41, 51) thereof is connected to the corresponding wall (4, 5), having a general shape which is corrugated about an axis (X-X) of the component (1, 2).

A main-body casing having an inlet port or an outlet port, a rotatable rotor shaft, and a rotor coupled to the rotor shaft are provided. A recess portion opened toward the inlet port is formed in the rotor and a fastening portion (a first shaft portion and a second shaft portion) of the rotor shaft is exposed to the recess portion. A cover portion, which is fastened to the fastening portion by a cover-portion fixing bolt and covers at least a part of the recess portion, is formed having a container shape and has a reinforcing portion located in a periphery of the fastening portion that prevents deflection by increasing rigidity. The cover portion also has a contact-pressure generating portion which is pressed by fastening to the fastening portion in a fastening direction and causes a contact pressure to be generated in the fastening direction.

A containment system for an engine includes an engine case having an inner perimeter. The containment system includes a containment ring nested within the inner perimeter of the engine case and integrally formed with the engine case along a first interface and a second interface. The containment ring includes a first leg opposite a second leg, and the first interface is defined between the first leg and the engine case. The containment system includes a first plurality of perforations defined at the first interface, and the first leg of the containment ring is frangible along the first plurality of perforations to at least partially release the containment ring to protect the engine case during a containment event.

A turbomachinery stator apparatus includes: a compressor casing including a casing wall defining an arcuate flowpath surface and an opposed backside surface, the flowpath surface defining at least two spaced-apart rotor lands; and a stator vane row of stator vanes disposed inside the compressor casing; wherein the casing wall includes at least one hollow structure; and wherein the casing wall is a single monolithic whole, wherein the stator vanes are integrally formed as part of the monolithic whole.

A rotor disk for a gas turbine engine includes a disk body having a central bore extending therethrough. The disk body includes a bore body that extends around the central bore, a web that extends radially outward from the bore body having decreased thickness relative to the bore body and a peripheral rim that is located at an outer end of the web. The peripheral rim includes blade mounting structures for engaging complimentary mounting structures of rotor blades. The bore body has a bore cavity that extends continuously through the bore body and about an entire periphery of the central bore. The bore cavity has a central axis that forms a circle about the central bore.

The present application provides a stator blade, a compressor structure and a compressor. The stator blade comprises a blade body, wherein a cavity is formed inside the blade body, and a gas supply hole is formed on the blade body. The present application forms a jet on the suction surface of the stator blade by supplemented gas, thereby blowing off the low-speed low-energy region formed by the suction surface, reducing the gas flow mixing loss caused by the supplemented gas, thereby improving the aerodynamic efficiency of the centrifugal compressor.

An air-sealing device intended to be inserted between an aircraft dual-flow turbine engine casing element and a nacelle element, the sealing device including an attachment tab at the end of which is located a sealing portion having an outer surface intended to be contacted by the casing element and the nacelle element, and an inner surface defining a cavity. The inner surface defines at least one protuberance extending inside the cavity.

An electric motor assembly includes an electric motor, a fan assembly coupled to the electric motor and configured to rotate therewith about an axis. The fan assembly includes a hub including a cylindrical portion having an inlet end and an outlet end, the hub further including an inlet surface coupled to the inlet end. The fan assembly also includes a plurality of blades coupled to an outer periphery of the cylindrical portion, wherein the inlet surface is tapered to direct an inlet airflow toward the plurality of blades.

An airfoil for a turbine engine includes an array of pins positioned in an internal cavity of the airfoil, such that cooling channels are defined in the interspaces between adjacent pins. Each pin extends lengthwise from a first airfoil wall to a second airfoil wall and is connected thereto at a first intersection and at a second intersection respectively. The pin has a first cross-sectional shape at a respective intersection and a second cross-sectional shape at an intermediate plane located between the first and second intersections. The first cross-sectional shape includes a closed shape defined by relatively sharp corners and the second cross-sectional shape includes a closed shape defined by relatively rounded corners. A cross-sectional area of the pin at the intermediate plane is greater than a cross-sectional area of the pin at the respective intersection.

Composite hollow blade and an associated method of forming the composite hollow blade are disclosed. The method includes forming a core by fabricating a grid core structure based on a plurality of design parameters, where the grid core includes a plurality of first reinforcing components disposed in a first curable matrix material. The method further includes forming an outer layer including a plurality of second reinforcing components disposed in a second curable matrix material. Further, the method includes coupling the core to the outer layer and curing the core and the outer layer to form the composite hollow blade.